Gas turbine engine with vane having a cooling inlet

ABSTRACT

An apparatus and method of cooling a hot portion of a gas turbine engine, such as a multi-stage compressor of a gas turbine engine, by reducing an operating air temperature in a space between a seal and a blade post of adjacent stages by routing cooling air through an inlet in the vane, passing the routed cooling air through the vane, and emitting the routed cooling air into the space.

CROSS REFERENCE TO RELATED APPLICATION

The present application is a continuation of U.S. patent applicationSer. No. 14/941,995, filed Nov. 16, 2015, and now allowed, which isincorporated herein by reference in its entirety.

BACKGROUND

Turbine engines, and particularly gas or combustion turbine engines, arerotary engines that extract energy from a flow of combusted gasespassing through the engine in a series of compressor stages, whichinclude pairs of rotating blades and stationary vanes, through acombustor, and then onto a multitude of turbine blades. In thecompressor stages, the blades are supported by posts protruding from therotor while the vanes are mounted to stator disks. Gas turbine engineshave been used for land and nautical locomotion and power generation,but are most commonly used for aeronautical applications such as forairplanes, including helicopters. In airplanes, gas turbine engines areused for propulsion of the aircraft.

Gas turbine engines for aircraft are designed to operate at hightemperatures to maximize engine thrust, so cooling of certain enginecomponents, such as the rotor post is necessary during operation.Typically, cooling is accomplished by ducting cooler air from the highand/or low pressure compressors to the engine components which requirecooling.

In adjacent compressor stages, there is a tendency for the pressureacross the adjacent stages to want to back flow through a seal with thevanes, leading to additional heating of the rotor post of an upstreamcompressor stage, which, under the certain thermal conditions, can leadto the temperature at the upstream rotor post exceeding its creeptemperature resulting unwanted creeping of the rotor post. This isespecially true for the most rearward or aft compressor stage, which issubject to the greatest temperature.

BRIEF DESCRIPTION

In one aspect the disclosure herein relates to a compressor for a gasturbine engine comprising an outer casing having circumferentiallyspaced vanes arranged in axially spaced groups of vanes, wherein eachvane comprises a pressure side and a suction side and extends axiallybetween a leading edge and a trailing edge; and a rotor located withinthe outer casing and having circumferentially spaced blades arranged inaxially spaced groups of blades in alternating axially arrangement withthe groups of vanes to define axially arranged pairs of vanes and blade,with each pair forming a compressor stage; the compressor stages havinga circumferential seal extending between the rotor and the vanes tofluidly seal axially adjacent compressor stages; and a cooling aircircuit passing through the vanes and having an elongated inlet locatedon one of the pressure side or suction side of the vanes and an outletat the rotor upstream of a corresponding seal for the vanes.

In another aspect the disclosure herein relates to a method of cooling amulti-stage compressor of a gas turbine engine, the method comprisingrouting compressor air through a vane having a pressure side and asuction side and located in of one of the stages by receiving thecompressor air through an elongated inlet located on one of the pressureside or suction side, passing the routed compressor air through thevane, and emitting the routed compressor into a space between the vaneand a blade of at least one of an upstream stage and downstream stage ofthe compressor.

In yet another aspect the disclosure herein relates to a vane assemblyfor a compressor of a gas turbine engine comprising a vane comprising apressure side and a suction side and having a leading edge and atrailing edge and a span extending from a root to a tip; a seal locatedon the root; and a cooling air circuit passing through the vane andhaving an elongated inlet located on one of the pressure side or suctionside of the vane and an outlet at a rotor, with the outlet located atleast one of upstream or downstream of the seal.

In yet another aspect the disclosure herein relates to a method ofcooling a multi-stage compressor of a gas turbine engine, the methodcomprising reducing an operating air temperature in a space between aseal and a blade post of adjacent stages at least 50 degrees Fahrenheitas compared to without the cooling, by routing compressor air through avane having a pressure side and a suction side and located in one of thestages by receiving the compressor air through an elongated inletlocated on one of the pressure side or suction side, passing the routedcompressor air through the vane, and emitting the routed compressor intothe space, which is upstream of the vane.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic, sectional view of a gas turbine engine accordingto an embodiment of the disclosure.

FIG. 2 is an enlarged, schematic cross-sectional partial view of aportion of the compressor in FIG. 1 and specifically including a vanebetween axially adjacent blades.

FIG. 3 is a perspective view of a vane suitable for use as the vane inFIGS. 1 and 2.

FIG. 4 is an enlarged, schematic sectional view of a lower portion ofthe vane of FIGS. 1 and 2.

FIG. 5 is a flow chart illustrating a method for reducing the operatingair temperature in a stage.

DETAILED DESCRIPTION

The described embodiments of the present disclosure are directed tosystems, methods, and other devices related to routing air flow in aturbine engine. For purposes of illustration, the present disclosurewill be described with respect to an aircraft gas turbine engine. Itwill be understood, however, that the disclosure is not so limited andmay have general applicability in non-aircraft applications, such asother mobile applications and non-mobile industrial, commercial, andresidential applications.

FIG. 1 is a schematic cross-sectional diagram of a gas turbine engine 10for an aircraft. The engine 10 has a generally longitudinally extendingaxis or centerline 12 extending forward 14 to aft 16. The engine 10includes, in downstream serial flow relationship, a fan section 18including a fan 20, a compressor section 22 including a booster or lowpressure (LP) compressor 24 and a high pressure (HP) compressor 26, acombustion section 28 including a combustor 30, a turbine section 32including a HP turbine 34, and a LP turbine 36, and an exhaust section38.

The fan section 18 includes a fan casing 40 surrounding the fan 20. Thefan 20 includes a plurality of fan blades 42 disposed radially about thecenterline 12. The HP compressor 26, the combustor 30, and the HPturbine 34 form a core 44 of the engine 10, which generates combustiongases. The core 44 is surrounded by an outer casing 46, which can becoupled with the fan casing 40.

A HP shaft or spool 48 disposed coaxially about the centerline 12 of theengine 10 drivingly connects the HP turbine 34 to the HP compressor 26.A LP shaft or spool 50, which is disposed coaxially about the centerline12 of the engine 10 within the larger diameter annular HP spool 48,drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20.The portions of the engine 10 mounted to and rotating with either orboth of the spools 48, 50 are also referred to individually orcollectively as a rotor 51.

The LP compressor 24 and the HP compressor 26 respectively include aplurality of compressor stages 52, 54, in which a set of compressorblades 58 rotate relative to a corresponding set of static compressorvanes 60, 62 (also called a nozzle) to compress or pressurize the streamof fluid passing through the stage. In a single compressor stage 52, 54,multiple compressor blades 56, 58 can be provided in a ring and canextend radially outwardly relative to the centerline 12, from a bladeplatform to a blade tip, while the corresponding static compressor vanes60, 62 are positioned downstream of and adjacent to the rotating blades56, 58. It is noted that the number of blades, vanes, and compressorstages shown in FIG. 1 were selected for illustrative purposes only, andthat other numbers are possible. The blades 56, 58 for a stage of thecompressor can be mounted to a disk 53, which is mounted to thecorresponding one of the HP and LP spools 48, 50, with each stage havingits own disk. The vanes 60, 62 are mounted to the core casing 46 in acircumferential arrangement about the rotor 51.

The HP turbine 34 and the LP turbine 36 respectively include a pluralityof turbine stages 64, 66, in which a set of turbine blades 68, 70 arerotated relative to a corresponding set of static turbine vanes 72, 74(also called a nozzle) to extract energy from the stream of fluidpassing through the stage. In a single turbine stage 64, 66, multipleturbine blades 68, 70 can be provided in a ring and can extend radiallyoutwardly relative to the centerline 12, from a blade platform to ablade tip, while the corresponding static turbine vanes 72, 74 arepositioned upstream of and adjacent to the rotating blades 68, 70. It isnoted that the number of blades, vanes, and turbine stages shown in FIG.1 were selected for illustrative purposes only, and that other numbersare possible.

In operation, the rotating fan 20 supplies ambient air to the LPcompressor 24, which then supplies pressurized ambient air to the HPcompressor 26, which further pressurizes the ambient air. Thepressurized air from the HP compressor 26 is mixed with fuel in thecombustor 30 and ignited, thereby generating combustion gases. Some workis extracted from these gases by the HP turbine 34, which drives the HPcompressor 26. The combustion gases are discharged into the LP turbine36, which extracts additional work to drive the LP compressor 24, andthe exhaust gas is ultimately discharged from the engine 10 via theexhaust section 38. The driving of the LP turbine 36 drives the LP spool50 to rotate the fan 20 and the LP compressor 24.

Some of the ambient air supplied by the fan 20 can bypass the enginecore 44 and be used for cooling of portions, especially hot portions, ofthe engine 10, and/or used to cool or power other aspects of theaircraft. In the context of a turbine engine, the hot portions of theengine are normally downstream of the combustor 30, especially theturbine section 32, with the HP turbine 34 being the hottest portion asit is directly downstream of the combustion section 28. Other sources ofcooling fluid can be, but is not limited to, fluid discharged from theLP compressor 24 or the HP compressor 26.

Referring to FIG. 2, a portion of the compressor section 22 is describedin greater detail and which includes axially adjacent blades 58 fromadjacent compressor stages 52, 54, with the intervening vane 62. Theblades 58 are mounted to a post 112 extending from a disk 53 of therotor 51. The blades 58 include a dovetail 76 that is received within aslot 114 in the post 112.

The vane 62 extend between inner and outer rings 100, 104. Each vane 62spans radially from a root 106, at the inner ring 100, to a tip 108, atthe outer ring 104. The vane 62 includes a leading edge 134 located onthe upstream side of the vane 62 and a trailing edge 136 on thedownstream side. Each vane further includes a pressure side 122 and asuction side 135 (FIG. 3). The vane 62 is mounted at the root 106 to theinner ring 100 and at the tip to the outer ring.

The inner ring 100 comprises an upper portion 118 and a lower portion120 which together form a circumferential channel 102 about the rotor51. The upper portion 118 provides a base 116 to which the root 106 ofthe vane 62 is mounted. The circumferential channel 102 provides an airconduit 110 circumferentially about the rotor 51.

A seal element, referred to as a seal 128 comprising a honeycomb element124 and annular fingers 126 seals the inner ring 100 relative to therotor 51. The honeycomb element 124 is mounted to the lower portion 120of the inner ring 100 and the annular fingers 126 project from the rotor51. The abutting of the fingers 126 with the honeycomb element 124 formsthe seal 128 to seal adjacent compressor stages 52, 54 thereby keepingcompressed air from a downstream compressor stage (relatively higherpressure) from backflowing to an upstream compressor stage (relativelylower pressure).

A seal cavity in the form of a space 130 is located upstream of the seal128 of the vane 62, radially inward of the blade 58 and between the seal128 and the post 112. As illustrated by the dashed arrow, under certainconditions backflow air 132 from a downstream compressor stage, whichhas higher temperatures due to being compressed and closer proximity tothe combustor 30, can collect in the space 130 causing the air in thespace 130 to increase in temperature. The compressor air flowing throughthe compressor stages 52,54 tends to flow over the space 130 and retardsthe removal of the backflow air 132, which leads to the generation of ahot spot in front of the post 112, and which leads to a heating of thepost 112.

A cooling air circuit 138 is provided to direct the compressor air intothe space 130 to effect a cooling of the space 130. The cooling aircircuit 138 has an inlet 140 on the vane 62 which opens to a vanechannel 142 within the vane 62. The vane channel 142 can be a hollowinterior of the vane 62 or a dedicated channel formed in the vane, suchas by placing an insert within the hollow interior. The vane channel 142is fluidly connected to the circumferential channel 102 of the innerring 100. An outlet 144 on the lower portion 120 of the inner ring 100is fluidly coupled to the circumferential channel 102 and located at theupstream of the inner ring 100 such that the outlet 144 opens to thespace 130. The outlet is not limited to upstream of the seal 128 and canbe located downstream of the inner ring 100.

The inlet 140 to the cooling air circuit 138 is best seen in FIG. 3,which is a perspective view of an exemplary 3-D vane 62. The inlet 140is elongated in the flow direction and extends generally between theleading edge 134 to the trailing edge 136 in a generally axialdirection. The inlet 140 is located on the pressure side 122 of the vane62, which will aid in the natural flowing of the compressor air into theinlet 140. However, while not as ideal, the inlet 140 can be located onthe suction side 135 of the vane 62. While the inlet 140 is illustratedas a simple opening, it can have a complex geometry, especially one thatdoes not follow the contour of the vane 62. A scoop or similar structurecan be provided for the inlet 140 to aid in directing the compressor airinto the inlet 140, which would be especially useful if the inlet 140 islocated on the suction side 135.

While the inlet 140 can be located anywhere along the span of the vane62, it is most advantageous for the inlet 140 to be located where thecoolest air is found. For most vane geometries, the coolest air islocated on the pressure side 122 of the vane 62, and at the mid-spanarea of the vane 62. The routed air follows a pathway 146 and isdeposited in the space 130. The pathway 146 begins on the leading edge134 of the vane 62 and passes through the cooling air circuit 138 intothe seal cavity space 130. The cooling air circuit 138 is provided in atleast some of the vanes 62 in the most downstream compressor stage 54 ofthe gas turbine engine 10.

The outlet 144 to the cooling air circuit 138 is best seen in FIG. 4,which is an enlarged schematic sectional view of the lower portion ofthe vane 62 taken from FIG. 2. While only one outlet 144 is shown, therecan be multiple outlets for each blade post 112. The outlets 144 arelocated such the air emitted from the outlet flows into the space 130and can impinge on the post 112.

It should be noted that while only one blade post 112 and one vane 62are described, that the description applies to all blade posts 112 andvanes 62 that are spaced circumferentially around the engine centerline12. It should also be noted that while the blade posts 112 and vanes 62are described for the compressor, that cooling air circuit 138 hassimilar applicability in other areas of the engine 10.

In operation, as shown in FIG. 2, backflow air 132 can develop betweenadjacent compressor stages, when the higher pressure air of thedownstream compressor passes upstream through the circumferential seal128. In essence, the higher pressure air in the downstream stage, whichis higher in temperature than the upstream stage, is suctioned by thepressure differential toward the upstream stage. The additional heatfrom the downstream stage heats the blade post 112. Under the rightconditions, the blade post can be heated to the point where it exceedsthe creep temperature and the blade post creeps radially. If the creepis greater enough, the blade 58 can rub the casing.

Cooling air from the cooling air circuit 138 is provided from the normalflow through the compressor stage along pathway 146 to the space 130 tocool the blade post 112. The cooling air supplied through the pathway146 drops the air temperature of the space 130 below a creep temperatureof the blade post 112. The cooling air circuit 138 reduces the airtemperature of the space 130 between the seal 128 and the blade post 112by at least 50 degrees Fahrenheit as compared to an air temperaturewithout cooling by routing compressor air through a cooling air circuit138.

FIG. 5 illustrates a method 300 of implementing the apparatus describedabove for cooling a multi-stage compressor of a gas turbine engine. Themethod 300 includes routing 302 compressor air which can follow thepathway 146 of FIG. 2 through the inlet 140 in the vane 62 of one of thecompressor stages 52, 54 in which the routed compressor air is drawnfrom a mid-span area of the vane 62 on the pressure side 122 of the vane62. The method further comprises passing 304 the routed compressor airthrough the vane 62, and finally emitting 306 the routed compressor intoa space 130 between the vane 62 and a blade 58 of at least one of anupstream stage 308 and downstream stage 310 of the compressor 24, 26.This method 300 reduces an operating air temperature in the space 130between the seal 128 and the blade post 112 of adjacent stages to belowthe creep temperature of the blade post 112. This method reduces the airtemperature at least 50 degrees Fahrenheit as compared to without thecooling.

This written description uses examples to illustrate the disclosure,including the best mode, and also to enable any person skilled in theart to practice the disclosure, including making and using any devicesor systems and performing any incorporated methods. The patentable scopeof the disclosure is defined by the claims, and may include otherexamples that occur to those skilled in the art. Such other examples areintended to be within the scope of the claims if they have structuralelements that do not differ from the literal language of the claims, orif they include equivalent structural elements with insubstantialdifferences from the literal languages of the claims.

What is claimed is:
 1. A compressor for a gas turbine engine, thecompressor comprising: an outer casing; a set of inner and outer ringsdisposed within the outer casing and having circumferentially spacedvanes arranged in axially spaced groups of vanes, wherein each vanecomprises a pressure side and a suction side and extends axially betweena leading edge and a trailing edge; and a rotor located within the outercasing and having circumferentially spaced blades arranged in axiallyspaced groups of blades in alternating axially arrangement with thegroups of vanes to define axially arranged pairs of vanes and blades,with each pair forming a compressor stage; the compressor stages havinga circumferential seal extending between the rotor and the vanes tofluidly seal axially adjacent compressor stages; and a cooling aircircuit passing through the vanes and having an elongated inlet locatedon one of the pressure side or suction side of the vanes and an outletat the rotor upstream of a corresponding seal for the vanes, wherein theinlet is located along a span of the respective vane in an area ofcoolest air flow over the vane.
 2. The compressor of claim 1, whereinthe inlet is located in a mid-span area of the vane.
 3. The compressorof claim 2, wherein the inlet is elongated in a flow direction.
 4. Thecompressor of claim 3, wherein the inlet is located on the pressure sideof the vane.
 5. The compressor of claim 4, wherein the inlet comprises ascoop.
 6. The compressor of claim 1, wherein the cooling air circuit isprovided in at least some of the vanes in the most downstream compressorstage.
 7. The compressor of claim 1, further comprising a first innerring of the set of inner rings, the first inner ring located within thecasing and supporting the vanes of the respective compressor stage at aroot of the corresponding vane and the first inner ring defines acircumferential channel forming part of the cooling air circuit.
 8. Thecompressor of claim 7, wherein the outlet of the cooling air circuit isformed in the ring.
 9. The compressor of claim 8, wherein the sealcomprises a honeycomb element mounted to the ring and fingers extendingfrom the rotor and abutting the honeycomb element.
 10. The compressor ofclaim 1, wherein the rotor comprises posts and the outlet emits thecooling air toward the post upstream of the vane.
 11. The compressor ofclaim 10, wherein a space between the posts of one compressor stage andseal for a downstream compressor stage define a seal cavity and theoutlet emits cooling air into the seal cavity.
 12. A method of cooling amulti-stage compressor of a gas turbine engine, the method comprisingrouting compressor air through a vane having a pressure side and asuction side and located in of one of the stages by receiving thecompressor air through an elongated inlet located on one of the pressureside or suction side, passing the routed compressor air through thevane, and emitting the routed compressor into a space between the vaneand a blade of at least one of an upstream stage and downstream stage ofthe compressor, wherein the inlet is located along a span of the vane inan area of coolest air flow over the vane.
 13. The method of claim 12,wherein the space is upstream of a seal for the vane.
 14. The method ofclaim 13, wherein the space is radially inward of the blade.
 15. Themethod of claim 14, wherein the space is between the seal and a postmounting the blade.
 16. The method of claim 12, wherein the routedcompressor air is drawn from a mid-span area of the vane.
 17. The methodof claim 12, wherein the routed compressor air is drawn from thepressure side of the vane.
 18. A vane assembly for a compressor of a gasturbine engine comprising: a vane comprising a pressure side and asuction side and having a leading edge and a trailing edge and a spanextending from a root to a tip; a seal located on the root; and acooling air circuit passing through the vane and having an elongatedinlet located on one of the pressure side or suction side of the vaneand an outlet at a rotor, with the outlet located at least one ofupstream or downstream of the seal, wherein the inlet is located along aspan of the vane in an area of coolest air flow over the vane.
 19. Thevane assembly of claim 18 wherein the inlet is located on one of amid-span area of the vane or the pressure side of the vane.